The disclosed embodiments generally relate to one or more structures for cooling an airfoil. More particularly, but not by way of limitation, present embodiments relate to trench cooling of airfoils including, but not limited to, a nozzle.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, a turbine, and an outlet/exhaust at the aft end of the engine. It will be readily apparent from those skilled in the art that additional components may also be included in the gas turbine engine, such as, for example, low-pressure and high-pressure compressors, and high-pressure and low-pressure turbines. This, however, is not an exhaustive list. The gas turbine engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy wherein each set of stator vanes turns and accelerates the combustion gases to engage an adjacent row of rotating turbine blades.
In the formation of components for aircraft and aircraft engines, such as for non-limiting example, turbine structures, blades, vanes and shrouds, various components are insulated from heat by thermal barrier coatings (“TBCs”), but most rely on various types of air-cooling to reduce or control temperature. For example, film cooling injects a thin blanket of cool air over one or more surfaces of the components, while effusion cooling pushes cool air through a lattice formed of closely spaced, discrete pores, or holes, in the component.
The cooling film holes are utilized in order to attain temperatures that are within limits of the part so that the part or component does not deteriorate or become damaged in the high temperature, pressure and stress environment of gas turbine engines. These cooling film holes receive bypass or cooling air within the aircraft engine to pass through the parts or components and provide the cooling necessary for operation in the extreme conditions. Current cooling film holes are formed by machining the cooling film holes into the component after the component has been cast. This adds cost and time to the process of forming the components. Additionally, the current technology being shaped diffuser holes are highly dependent on geometry of the diffuser, its feed hole and the cast wall thickness.
Reliable and accurate inspection of diffuser holes is currently difficult and sometimes requires destructive evaluation or cutup of the component. This is undesirable.
As may be seen by the foregoing, it would be desirable to overcome these and other difficulties with cooling systems of gas turbine engine components.